1. Field of the Invention
The present invention relates generally to fluid reaction surfaces, and more specifically to turbine rotor blade with a trailing edge cooling circuit.
2. Description of the Related Art Including Information Disclosed Under 37 CFR 1.97 and 1.98
A gas turbine engine includes a turbine section with multiple stages of stator vanes and rotor blades that are exposed to a high temperature gas flow produced in the combustor section by burning a fuel. The engine efficiency can be increased by passing a higher gas flow temperature into the turbine section. The material properties of the first stage stator vanes and rotor blades establish a maximum temperature for the turbine section.
These high temperature turbine airfoils are provided with complex internal cooling circuit to provide cooling for the airfoils to extend the operating temperature of these airfoils beyond the material characteristic temperature limits. Hot spots can appear on the airfoils due to uneven exposure to the hot gas flow through the turbine and to uneven cooling provided by the convection and film cooling circuits. Especially in an industrial gas turbine—where engine part life is a major design factor—a large rotor blade will produce high levels of stress in the lower portions of the blade closer to the platform. Higher amounts of cooling air required for the portions of the blade that would produce high levels of creep. In other words, more cooling is required in the lower sections of the rotor blade because of the high stress levels that occur due to the mass of the rotating blade. The blade section above the root will tend to pull on the blade section near the root due to centrifugal forces that develop during rotation of the blade. Because the rotor blade is exposed to an extremely high temperature, and that the blade material becomes weaker as the temperature of the metal rises, without adequate cooling at the lower section the rotor blade could have problems with excessive creep. This will shorten the life of the blade and require premature engine over hall to fix damaged blades.
Prior art turbine blade have cooling holes drilled into the trailing edge region of the blade that connect to an internal cooling air supply channel formed into the turbine blade. Cooling air flows upward in the cooling air supply passage and bleeds off into the row of cooling holes to provide cooling for the trailing edge region. This single pass axial flow cooling circuit of the prior art design provides very little cooling for the trailing edge region because the flow path for the cooling air is very short. U.S. Pat. No. 5,387,085 issued to Thomas, Jr et al on Feb. 7, 1995 and entitled TURBINE BLADE COMPOSITE COOLING CIRCUIT discloses this blade trailing edge cooling circuit. Also, the lower reaches of the blade have low levels of cooling while the upper reaches (near the tip) have too much cooling. Creep is a major problem in the lower reaches of the blade and decreases in the direction of the blade tip.
U.S. Pat. No. 6,491,496 B2 issued to Starkweather on Dec. 10, 2002 and entitled TURBINE AIRFOIL WITH METERING PLATES FOR REFRESHER HOLES shows a rotor blade with the cooling air supply channel following a serpentine flow path before the cooling air is bled off into the trailing edge cooling holes. The cooling air supply path to the trailing edge cooling holes is longer and therefore the cooling air gains more heat prior to discharging out through the exit holes along the trailing edge.
Another prior art device, U.S. Pat. No. 6,139,269 issued to Liang on Oct. 31, 2000 and entitled TURBINE BLADE WITH MULTI-PASS COOLING AND COOLING AIR ADDITION shows the trailing edge region being cooled by a circuit that used multiple impingement cooling in the trailing edge region. This design improves the trailing edge region cooling capability over the above cited prior art cooling circuits.
U.S. Pat. No. 6,099,252 issued to Manning et al on Aug. 8, 2000 and entitled AXIAL SERPENTINE COOLED AIRFOIL discloses a turbine blade with the trailing edge region cooled by an axial serpentine cooling circuit having a plurality of serpentine circuits stacked in a radial row along the airfoil trailing edge. These stacked serpentine circuits form a plurality of parallel cooling circuits connected to the cooling supply channel. The flow path of the cooling air through the trailing edge region is increased over the above cited prior art circuits and thus the cooling ability of the Manning et al circuit is increased.
It is therefore an object of the present invention to provide for a cooling circuit in an airfoil trailing edge region that will reduce the metal temperature and thus reduce the cooling flow requirement over the above cited prior art trailing edge cooling circuits.